(1) Field of the Invention
The invention relates to a fatigue management system and a method of operating a fatigue management system.
(2) Description of Related Art
Components of helicopters or fixed wing aircrafts have a limit service life to maintain operational safety, said service life limit (SLL) or component retirement time (CRT) being evaluated during load classification flights of e.g. a model of the helicopter. Said service life limit is adapted to several mission profiles while taking into account redundancy and safety assumptions. For customers having helicopters or fixed wing aircrafts not flying any of these mission profiles any service life limit of components may not be exhausted even though due to the standards of the evaluation said components have to be serviced.
The document DE 4336588 A1 discloses a method for determining the individual life of an aircraft by means of a plurality of neural networks for flight attitudes, centres of gravity and weights and their sequences in time, which are linked to one another and whose results are stored in an on-board black box, as well as being entered in an external ground station for detailed determination of the current individual life and the individual maintenance intervals.
The document U.S. Pat. No. 6,480,792 B1 discloses a fatigue monitoring system and method, in which a stream of data relating to the stresses experienced at a plurality of locations over the structure during operation is applied to a ground based neural network trained to remove data stream values deemed to be in error. The stresses are either signalled by a plurality of sensors disposed at different locations in said structure or said stresses are calculated by comparison with a large number of templates, said templates being derived from finite element analysis and the results of ground based airframe fatigue tests. The data from the neural network is then processed to determine the fatigue life.
The document U.S. Pat. No. 4,336,595 A discloses an electronic device which determines the fatigue life of a structure and its crack growth characteristics when subjected to repeated loading by using the signal from an electrical resistance type strain gage secured to the structure to provide the input to a processor which monitors cyclic excursions, calculates the fatigue and fracture structural damage from those excursions, stores the cumulative damage, and displays that damage on command
The document US2010100338 A1 discloses a monitoring system, which may include a structural component configured to undergo mechanical loading and a wireless node attached to the structural component. The node may include a strain sensing device configured to measure strain experienced by the structural component at the location of the node. The node may also include a processor configured to predict, based on the strain measurements, fatigue life of the structural component
The document US2008167833 A1 discloses methods and systems for structural and component health monitoring. A system includes a plurality of sensor systems positioned about an object to be monitored and a processing system communicatively coupled to at least one of said plurality of sensor systems wherein the processing system includes an ontology and reasoning module configured to model the object to be monitored, reason about the received sensor data associated with the object to be monitored and reason about the relationships between the received sensor data to fuse the data into contextual information for the overall object to be monitored and a contextual analyzer configured to transmit the received sensor data to said ontology and reasoning module and to store the information into a contextual information database.
The document US2009306909 A1 discloses a method for the evaluation of measured values for the recognition of defect conditions due to material fatigue on aircraft parts, in which strain sensors are applied on the critical locations of an aircraft or the parts thereof, wherein the measured values of the strain sensors at different loading conditions are detected, amplified and stored through several measuring circuits and from which an evaluating apparatus derives, signals or indicates a material fatigue by comparison of current measured values with previous measured values. Critical aircraft parts are impinged on with a plurality of prescribed loading conditions by a plurality of loading elements. The strain effect caused thereby is detected by a plurality of measuring circuits, and the evaluating apparatus forms, for at least each loading condition and each measuring circuit, an allocated reference value and a permissible limit value range, which is subsequently coupled with the current measured values in such a manner so that the exceeding of the limit value range represents a material fatigue manifestation.